Gas turbine engine with a fluid conduit system and a method of operating the same

ABSTRACT

A method of operating a gas turbine engine comprising: extracting a flow of air from a compressor section of the gas turbine engine into a first conduit; flowing the extracted flow of air through the first conduit to a first location at a turbine section of the turbine section, wherein a second conduit is in fluid communication with the turbine section at a second location; flowing a heat transfer fluid to a first heat exchanger positioned in thermal communication with the flow of air through the first conduit, the heat transfer fluid in thermal communication with the extracted flow of air through the first conduit via the first heat exchanger; and modulating, via a flow control device, a portion of the flow of air extracted from the first conduit to the second conduit downstream of the first heat exchanger.

GOVERNMENT SPONSORED RESEARCH

The project leading to this application has received funding from theEuropean Union Clean Sky 2 research and innovation program under grantagreement No. CS2-ENG-GAM-2014-2015-01.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority to Polish Patent Application No.P.439448, filed Nov. 5, 2021, which is a non-provisional application,and wherein the above application is hereby incorporated by reference inits entirety.

FIELD

The present subject matter relates particularly to fluid conduits forgas turbine engines, such as for clearance control structures of turbinesections of gas turbine engines.

BACKGROUND

Casings for gas turbine engines, such as turbine section casingssurrounding turbine section rotors, generally require separable flangesand assembled casing and manifold portions due to internally andexternally mounted components. Such components generally includebrackets or hangers for turbine shrouds, or flanges for multiplecasings. Additionally, since turbine casings surround turbine rotors,excessive deformation, thermal expansion or contraction, or bowing mayresult in excessive rub and undesired contact with the turbine rotors,which can result in loss in performance or operability. Certain casingsmay include assemblies via separable flanges to limit deformation ordisplacement during engine operation and thermal cycling. However, theinventors of the present disclosure have found that such designs requireassembly and parts that add weight to the engine. Moreover, theinventors of the present disclosure have found that such designs mayfurther inhibit the inclusion or placement of thermal control structuresfor more effective clearance control.

As such, the inventors of the present disclosure have found that thereis a need for turbine casings that can overcome these limitations andprovide improved thermal control, improved engine efficiency, andreduced weight.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present disclosure, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is an exemplary schematic cross sectional view of an embodimentof a gas turbine engine in accordance with aspects of the presentdisclosure;

FIG. 2 is an exemplary schematic cross sectional view of an embodimentof a gas turbine engine in accordance with aspects of the presentdisclosure;

FIG. 3 is an exemplary schematic cross sectional view of an embodimentof a gas turbine engine in accordance with aspects of the presentdisclosure;

FIG. 4 is a schematic cross-sectional view of a portion of an embodimentof a gas turbine engine in accordance with aspects of the presentdisclosure;

FIG. 5 is a schematic cross-sectional view of a portion of an embodimentof a gas turbine engine in accordance with additional aspects of thepresent disclosure;

FIG. 6 is a perspective view of a portion of an embodiment of the gasturbine engine in accordance with aspects of the present disclosure;

FIGS. 7A-7B depict flowcharts outlining steps of a method for operatingan engine in accordance with aspects of the present disclosure;

FIGS. 8-11 are exemplary schematic cross-sectional views of embodimentsof a portion of a turbine section and casing in accordance with aspectsof the present disclosure;

FIG. 12 is an exemplary perspective view of an embodiment of a portionof a manifold of the turbine section in accordance with aspects of thepresent disclosure;

FIGS. 13A-13D are exemplary sectional views of an embodiment of themanifold provided in FIG. 12 ;

FIG. 14 is an exemplary schematic cross-sectional view of an embodimentof a portion of a turbine section and casing in accordance with aspectsof the present disclosure;

FIG. 15 is an exemplary perspective view of an embodiment of a portionof a manifold of the turbine section in accordance with aspects of thepresent disclosure;

FIG. 16 is an exemplary schematic cross sectional view of an embodimentof a portion of a turbine section and casing in accordance with aspectsof the present disclosure;

FIG. 17 is a detailed view of an exemplary schematic cross sectionalview of the embodiment of FIG. 16 in accordance with aspects of thepresent disclosure;

FIG. 18 is a top-down view of an exemplary embodiment of a plurality ofpins of the thermal control ring in accordance with aspects of thepresent disclosure;

FIG. 19 is an exemplary schematic of flows of air through the turbinesection and casing of FIG. 16 in accordance with aspects of the presentdisclosure;

FIG. 20 is a perspective view of a portion of the engine in accordancewith aspects of the present disclosure; and

FIG. 21 is a cross-sectional view of the embodiment of the engineprovided in FIG. 20 in accordance with aspects of the presentdisclosure.

Repeat use of reference characters in the present specification anddrawings is intended to represent the same or analogous features orelements of the present disclosure.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of thedisclosure, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the disclosure.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations. Additionally, unlessspecifically identified otherwise, all embodiments described hereinshould be considered exemplary.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 1, 2, 4,10, 15, or 20 percent margin.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

Pressure values, and ranges thereof, are in absolute pressuremeasurement (psia) or equivalent. Values and ranges of pressure providedherein may be converted to ranges in gauge pressure, or other pressureunits, or other units, measurements, or combinations thereof thatcorrespond to the values and/or ranges disclosed herein.

The term “overall power output” refers to a maximum rated power outputof an engine.

The term “operating envelope” refers to a cycle, mission, or set ofmaneuvers at which the engine may normally operate. In one embodiment, alanding-takeoff (LTO) cycle may define an operating envelope. The LTOcycle including one or more combinations of startup, idle, takeoff,cruise, and approach engine operating conditions may collectively definethe operating envelope. In various embodiments, the cruise conditiondefines a majority of the operating envelope, such as to define amajority of an operating time or duration of the engine operation. Incertain embodiments, the cruise condition is between approximately 55%and 75% of the operating envelope. Stated differently, the cruisecondition may define approximately 55% to approximately 75% of theduration of engine operation from startup to shutdown following approachoperating condition. In another embodiment, the cruise condition maydefine approximately 60% to approximately 70% of the duration of engineoperation.

The term “cruise operating condition” may further refer to mid-powerengine operating condition. The term “takeoff operating condition” mayrefer to a full power condition and “idle operating condition” may referto a low power condition, and “cruise operating condition” is a power orthrust condition therebetween. In some embodiments, the cruise conditioncorresponds to approximately 75% to approximately 90% of an overallpower output of the engine. In still certain embodiments, the cruisecondition corresponds to approximately 80% to 88% of the overall poweroutput of the engine.

A “third stream” as used herein means a non-primary air stream capableof increasing fluid energy to produce a minority of total propulsionsystem thrust. A pressure ratio of the third stream may be higher thanthat of the primary propulsion stream (e.g., a bypass or propellerdriven propulsion stream). The thrust may be produced through adedicated nozzle or through mixing of an airflow through the thirdstream with a primary propulsion stream or a core air stream, e.g., intoa common nozzle.

In certain exemplary embodiments an operating temperature of the airflowthrough the third stream may be less than a maximum compressor dischargetemperature for the engine, and more specifically may be less than 350degrees Fahrenheit (such as less than 300 degrees Fahrenheit, such asless than 250 degrees Fahrenheit, such as less than 200 degreesFahrenheit, and at least as great as an ambient temperature). In certainexemplary embodiments these operating temperatures may facilitate heattransfer to or from the airflow through the third stream and a separatefluid stream. Further, in certain exemplary embodiments, the airflowthrough the third stream may contribute less than 50% of the totalengine thrust (and at least, e.g., 2% of the total engine thrust) at atakeoff condition, or more particularly while operating at a ratedtakeoff power at sea level, static flight speed, 86 degree Fahrenheitambient temperature operating conditions.

Furthermore in certain exemplary embodiments, aspects of the airflowthrough the third stream (e.g., airstream, mixing, or exhaustproperties), and thereby the aforementioned exemplary percentcontribution to total thrust, may passively adjust during engineoperation or be modified purposefully through use of engine controlfeatures (such as fuel flow, electric machine power, variable stators,variable inlet guide vanes, valves, variable exhaust geometry, orfluidic features) to adjust or optimize overall system performanceacross a broad range of potential operating conditions.

The term “turbomachine” or “turbomachinery” refers to a machineincluding one or more compressors, a heat generating section (e.g., acombustion section), and one or more turbines that together generate atorque output.

The term “gas turbine engine” refers to an engine having a turbomachineas all or a portion of its power source. Example gas turbine enginesinclude turbofan engines, turboprop engines, turbojet engines,turboshaft engines, etc.

The term “combustion section” refers to any heat addition system for aturbomachine. For example, the term combustion section may refer to asection including one or more of a deflagrative combustion assembly, arotating detonation combustion assembly, a pulse detonation combustionassembly, or other appropriate heat addition assembly. In certainexample embodiments, the combustion section may include an annularcombustor, a can combustor, a cannular combustor, a trapped vortexcombustor (TVC), or other appropriate combustion system, or combinationsthereof.

The terms “low” and “high”, or their respective comparative degrees(e.g., -er, where applicable), when used with a compressor, a turbine, ashaft, or spool components, etc. each refer to relative speeds within anengine unless otherwise specified. For example, a “low turbine” or “lowspeed turbine” defines a component configured to operate at a rotationalspeed, such as a maximum allowable rotational speed, lower than a “highturbine” or “high speed turbine” at the engine.

The term “at,” as used herein to refer to a location of a first objectrelative to a second object (e.g., the first object located orpositioned at the second object) refers to the first object beingpositioned wholly or partially within the second object, the firstobject contacting the second object, or the first object beingpositioned closest to the second object (relative to any othersurrounding relevant components).

One or more components of the turbomachine engine described herein belowmay be manufactured or formed using any suitable process, such as anadditive manufacturing process, such as a 3-D printing process. The useof such a process may allow such component to be formed integrally, as asingle monolithic component, or as any suitable number ofsub-components. In particular, the additive manufacturing process mayallow such component to be integrally formed and include a variety offeatures not possible when using prior manufacturing methods. Forexample, the additive manufacturing methods described herein may allowfor the manufacture of passages, conduits, cavities, openings, casings,manifolds, double-walls, heat exchangers, or other components, orparticular positionings and integrations of such components, havingunique features, configurations, thicknesses, materials, densities,fluid passageways, headers, and mounting structures that may not havebeen possible or practical using prior manufacturing methods. Some ofthese features are described herein.

Suitable additive manufacturing techniques in accordance with thepresent disclosure include, for example, Fused Deposition Modeling(FDM), Selective Laser Sintering (SLS), 3D printing such as by inkjets,laser jets, and binder jets, Stereolithography (SLA), Direct SelectiveLaser Sintering (DSLS), Electron Beam Sintering (EBS), Electron BeamMelting (EBM), Laser Engineered Net Shaping (LENS), Laser Net ShapeManufacturing (LNSM), Direct Metal Deposition (DMD), Digital LightProcessing (DLP), Direct Selective Laser Melting (DSLM), Selective LaserMelting (SLM), Direct Metal Laser Melting (DMLM), and other knownprocesses.

Suitable powder materials for the manufacture of the structures providedherein as integral, unitary, structures include metallic alloy, polymer,or ceramic powders. Exemplary metallic powder materials are stainlesssteel alloys, cobalt-chrome, aluminum alloys, titanium alloys, nickelbased superalloys, and cobalt based superalloys. In addition, suitablealloys may include those that have been engineered to have goodoxidation resistance, known as “superalloys” which have acceptablestrength at the elevated temperatures of operation in a gas turbineengine, e.g. Hastelloy, Inconel alloys (e.g., IN 738, IN 792, IN 939),Rene alloys (e.g., Rene N4, Rene N5, Rene 80, Rene 142, Rene 195),Haynes alloys, Mar M, CM 247, CM 247 LC, C263, 718, X-850, ECY 768, 282,X45, PWA 1483 and CMSX (e.g. CMSX-4) single crystal alloys. Themanufactured objects of the present disclosure may be formed with one ormore selected crystalline microstructures, such as directionallysolidified (“DS”) or single-crystal (“SX”).

Embodiments of a gas turbine engine including an improved clearancecontrol system are provided. The engine reduces weight and tubes,manifolds, or conduits outside of an outer core casing or fan casing byreducing or eliminating air extracted from a fan bypass passage forcooling at a turbine section. Embodiments provided herein allow forengines without fan casings, such as open rotor engines or propfanengines, to have and operate improved clearance control, coolingsystems, or air systems for turbine sections and/or bearing assemblies.It should be appreciated that while such embodiments may be applied toturbofan engines including nacelles and fan casings, embodimentsprovided herein allow for engines without nacelles, fan casings, orother structures surrounding the fan section to receive air for turbinesection cooling, clearance control, or bearing assemblies.

The improved gas turbine engine provided herein may additionally, oralternatively, allow for lower-pressure and/or lower-temperature air tobe removed from the compressor section for cooling or clearance controlat the turbine section and bearing assembly. Certain clearance controlsystems may generally utilize high-energy air (i.e., high-pressureand/or high-temperature air), such as from aft stages of a high pressurecompressor, and mix with one or more other sources of air, such as fromother compressor stages or from the fan air stream. Such high-energy airreduces engine efficiency, such as by removing energy from thethermodynamic and combustion process, or by requiring greater reductionin heat load before the air is appropriate for cooling or clearancecontrol at the turbine section. Still further, certain clearance controlsystems may not be suitable for additionally providing air to a bearingassembly for cooling, buffer air, or other uses at the bearing assembly.

Another aspect of the disclosure is directed to an improved turbinecasing allowing for improved clearance control, cooling fluiddistribution, reduced weight, and improved engine efficiency.Embodiments of an engine, casing, and manifold provided herein includeintegral, unitary structures such as may be formed by additivemanufacturing processes that would not have heretofore been possible orpracticable. Embodiments depicted and described herein allow forimproved and advantageous positioning of thermal control rings forimproved clearance control response, improved formation and positioningof openings, passages, and conduits to allow for more efficient heattransfer fluid utilization and movement, and reduced weight, such as viaobviating flanges and sub-assemblies into integral components.Particular combinations of these features allow for improved heattransfer properties and reduced thermal gradients. Improved heattransfer properties particularly include a lower heat transfercoefficient at certain features, such as at the plurality of walls thatform thermal control rings as provided herein. Such improvements maymitigate or eliminate undesired or excessive deformation, ovalization,bowing, or other changes in casing geometry that may adversely affectdeflections or result in undesired contact to the turbine rotors.

Embodiments provided herein include, e.g., an integral, unitary highspeed turbine casing and turbine center frame or mid-turbine framepositioned downstream of the high speed turbine and upstream of a low-or intermediate-pressure turbine. Embodiments provided herein furtherinclude, e.g., an integral, unitary clearance control manifoldconfigured to provide heat transfer fluid to thermal control rings. Theintegral, unitary structures may further allow for improved positioningof the thermal control rings relative to the turbine rotors, such as toprovide improved clearance control across the turbine rotor assembly.

As used herein, the term “integral, unitary” as used to describe astructure refers to the structure being formed integrally of acontinuous material or group of materials with no seams, connectionsjoints, or the like. The integral, unitary structures described hereinmay be formed through additive manufacturing to have the describedstructure, or alternatively through a casting process, etc.

Referring now to the drawings, FIG. 1 is a schematic cross-sectionalview of an exemplary gas turbine engine 10 herein referred to as “engine10” as may incorporate various embodiments of the present disclosure.Particular embodiments of the engine 10 may be configured as a turbofan,turboprop, turboshaft, or propfan gas turbine engine, or one or more gasturbine engines configured as hybrid-electric gas turbine engines, orother gas turbine engine configuration.

As shown in FIG. 1 , the engine 10 has a longitudinal or axialcenterline axis 12 that extends therethrough for reference purposes. Ingeneral, the engine 10 may include a turbomachine 14 disposed downstreamfrom a fan section 16.

The engine 10 includes a compressor section 21 in serial flowarrangement with a turbine section 27. The turbomachine 14 may generallyinclude a substantially tubular outer casing 18 that defines an annularinlet 20. The outer casing 18 may be formed from multiple casings. Theouter casing 18 encases, in serial flow arrangement, the compressorsection 21, a combustion section 26, and the turbine section 27. In aparticular embodiment, the compressor section 21 includes a booster orlow speed compressor 22 and a high speed compressor 24. In a stillparticular embodiment, the turbine section 27 includes a first turbineassembly or high speed turbine 28 and a second turbine assembly or lowspeed turbine 30 (e.g., including vanes 116 and rotor blades 118). A jetexhaust nozzle section 32 is positioned downstream of the turbinesection 27. A high speed shaft or spool 34 drivingly connects the highspeed turbine 28 to the high speed compressor 24. A low speed shaft orspool 36 drivingly connects the low speed turbine 30 to the low speedcompressor 22. The low speed spool 36 may also be connected to a fanshaft or spool 38 of the fan section 16. In particular embodiments, thelow speed spool 36 may be connected directly to the fan spool 38 such asin a direct-drive configuration. In alternative configurations, as isdepicted in phantom in FIG. 1 , the low speed spool 36 may be connectedto the fan spool 38 via a gear assembly 37, such as to configure theengine 10 as an indirect-drive or geared-drive configuration allowingfor a higher or lower rotational speed of the fan spool 38 versus thelow speed spool 36. Such gear assemblies may be included between anysuitable shafts/spools within engine 10 as desired or required.

Although depicted and described as a two-spool engine including the highspeed spool 34 separately rotatable from the low speed spool 36, itshould be appreciated that the engine 10 may be configured as athree-spool engine including the high speed spool 34, the low speedspool 36, and a third spool or intermediate speed spool positioned inserial flow arrangement between the high speed spool 34 and the lowspeed spool 36. Accordingly, the compressor section 21 may include anintermediate speed compressor separately rotatable from the high speedcompressor 24 and the low speed compressor 22. Similarly, the turbinesection 27 may include a third turbine assembly or an intermediate speedturbine separately rotatable from the high speed turbine 28 and the lowspeed turbine 30. The intermediate speed compressor and the intermediatespeed turbine may together be coupled to form an intermediate speedspool fluidly between the high speed spool and the low speed spool.

It should further be appreciated that in certain embodiments the lowspeed turbine 30 or second turbine assembly described herein generallyrefers to a separately rotatable spool downstream of the high speedturbine or first turbine assembly. As such, the second turbine assemblymay include an intermediate speed turbine or a low speed turbinepositioned aft or downstream of the high speed turbine.

As shown in FIG. 1 , the fan section 16 includes one or moreaxially-spaced stages of a plurality of fan blades 40 that are coupledto and that extend radially outwardly from the fan spool 38. An annularfan casing or nacelle 42 circumferentially surrounds the fan section 16and/or at least a portion of the turbomachine 14. It should beappreciated that for the embodiment depicted the nacelle 42 is supportedrelative to the turbomachine 14 by a plurality ofcircumferentially-spaced outlet guide vanes 44.

A bypass airflow passage 48 is formed downstream of one or more stagesof the plurality of fan blades 40 and around an outer portion of theturbomachine 14. In a particular embodiment, such as depicted in FIG. 1, the bypass airflow passage 48 is defined at a downstream section 46 ofthe nacelle 42 (downstream of the outlet guide vanes 44) and between thenacelle 42 and the outer portion of the turbomachine 14.

However, in other embodiments, it should be appreciated that the lowspeed compressor 22 may form one or more stages of the fan section 16,such as depicted in FIG. 3 . As such, the bypass airflow passage 48 maygenerally include any flowpath downstream of one or more stages of theplurality of fan blades 40 or the low speed compressor 22 and bypassingor surrounding at least a portion of the high speed compressor 24, andhaving a flow of bypass air 177 therethrough provide thrust.Accordingly, certain embodiments of the engine 10 provided herein may beconfigured as a third stream or adaptive cycle engine having a pluralityof bypass airflow passages 48 downstream of one or more stages of theplurality of fan blades and/or the low speed compressor 22 and upstreamof at least a portion of the high speed compressor 24, with one or moreof which configured as a “third stream.”

The engine 10 includes a computing system 1210 configured to performoperations. The computing system 1210 is communicatively coupled to theturbomachine 14 and/or a starter motor (not depicted) to adjust,modulate, maintain, change, or articulate any one or more controlsurfaces to generate the flows of air, one or more embodiments of theflow of heat transfer fluid, and/or a liquid and/or gaseous fuel inaccordance with aspects of the present disclosure provided herein. Thecomputing system 1210 can generally correspond to any suitableprocessor-based device, including one or more computing devices. Certainembodiments of the computing system 1210 include a full authoritydigital engine controller (FADEC), a digital engine controller (DEC), orother appropriate computing device configured to operate the engine 10.

The computing system 1210 may include one or more processors 1212 andone or more associated memory devices 1214 configured to perform avariety of computer-implemented functions, such as steps of the methodsdescribed herein. As used herein, the term “processor” refers not onlyto integrated circuits referred to in the art as being included in acomputer, but also refers to a controller, microcontroller, amicrocomputer, a programmable logic controller (PLC), an applicationspecific integrated circuit (ASIC), a Field Programmable Gate Array(FPGA), and other programmable circuits. Additionally, the memory 1214can generally include memory element(s) including, but not limited to,computer readable medium (e.g., random access memory (RAM)), computerreadable non-volatile medium (e.g., flash memory), a compact disc-readonly memory (CD-ROM), a magneto-optical disk (MOD), a digital versatiledisc (DVD), non-transitory computer-readable media, and/or othersuitable memory elements or combinations thereof.

The computing system 1210 may include control logic 1216 stored in thememory 1214. The control logic 1216 may include computer-readableinstructions that, when executed by the one or more processors 1212,cause the one or more processors 1212 to perform operations, such asoutlined in one or more steps of the method 1000 provided further below.In still various embodiments, the memory 1214 may store charts, tables,functions, look ups, schedules etc. corresponding to the flows, orrates, pressures, or temperatures associated with the flows of air, heattransfer fluid, or fuel provided herein. The instructions can besoftware written in any suitable programming language or can beimplemented in hardware. Additionally, and/or alternatively, theinstructions can be executed in logically and/or virtually separatethreads on the processor(s).

The computing system 1210 may also include a communications interfacemodule 1230. In various embodiments, the communications interface module1230 can include associated electronic circuitry that is used to sendand receive data. As such, the communications interface module 1230 ofthe computing system 1210 can be used to receive data from one or morecontrol surfaces, sensors, measurement devices, or instrumentation, orcalculations or measurements corresponding to one or more portions ofthe engine 10 provided herein, and may execute one or more steps of themethod 1000 provided herein. The computing system(s) 1210 can alsoinclude a network interface used to communicate, for example, with theother components of engine 10. The network interface can include anysuitable components for interfacing with one or more network(s),including for example, transmitters, receivers, ports, controllers,antennas, and/or other suitable components.

It should be appreciated that the communications interface module 1230can be any combination of suitable wired and/or wireless communicationsinterfaces and, thus, can be communicatively coupled to one or morecomponents of the apparatus via a wired and/or wireless connection. Assuch, the computing system 1210 may obtain, determine, store, generate,transmit, or operate any one or more steps of the method describedherein via a distributed network. For instance, the network can includea SATCOM network, ACARS network, ARINC network, SITA network, AVICOMnetwork, a VHF network, a HF network, a Wi-Fi network, a WiMAX network,a gatelink network, etc.

Referring now to FIG. 2 , an exemplary embodiment of an open rotorconfiguration of the engine 10 depicted and described with regard toFIG. 1 is provided. The embodiment of the engine 10 provided in FIG. 2is configured substantially similarly as provided in FIG. 1 . However,in FIG. 2 , the open rotor configuration of the engine 10 does not havea fan casing or the nacelle 42 (depicted in FIG. 1 ) surrounding theplurality of fan blades 40. The bypass airflow passage 48 is formeddownstream of the plurality of fan blades 40, or particularly downstreamof the outlet guide vanes 44, and radially outward of the outer portionof the turbomachine 14.

Referring now to FIG. 3 , an exemplary embodiment of an open rotorconfiguration in accordance with FIG. 2 is provided. The embodimentprovided in FIG. 3 further includes a plurality of bypass airflowpassages 48 formed downstream of the plurality of fan blades 40, such asdescribed above. In the particular embodiment depicted, the engine 10includes a first bypass airflow passage 48A and a second bypass airflowpassage 48B. The second bypass airflow passage 48B is extended from alocation between the low speed compressor 22 and the high speedcompressor 24 to an exhaust to atmosphere (although in other embodimentsthe second bypass airflow passage 48B may extend to the first bypassairflow passage 48A). An articulating vane or door structure 43 may bepositioned at the second bypass airflow passage 48B. The door structure43 may include any appropriate type of actuatable wall, vane, door, orother structure configured to desirably alter a flow of air 172 receivedfrom a core gas flowpath 70 and allowed through the second bypassairflow passage 48B, such as depicted schematically via arrows 177. Thesecond bypass airflow passage 48B may be referred to as a third stream.

Although FIG. 3 depicts the engine 10 having a three-stream or adaptivecycle with an open rotor configuration, it should be appreciated thatthe adaptive cycle configuration may also include a nacelle surroundingthe fan section, such as depicted and described with regard to FIG. 1 .In such a manner, it should further be appreciated that although certainadvantages and benefits provided herein may provide benefits forturbofan engines having nacelles, embodiments and arrangements of thecomponents provided herein may overcome issues or challenges that areparticular to open rotor configurations.

Referring now to FIGS. 4-5 , enlarged cross-sectioned views of an engine10 configured in a similar manner as one or more of the exemplaryengines 10 depicted in FIGS. 1-3 are provided. FIGS. 4-5 depict walledconduits, manifolds, tubes, or other structures forming flowpathsconfigured to extract or receive a flow of air, depicted schematicallyvia arrows 91, from the compressor section 21 and provide the flow ofair 91 to the turbine section 27. The flow of air 91 provided to theturbine section 27 may be utilized for cooling blades, vanes, shrouds,or other portions of the turbine section 27. In certain embodiments, theturbine section 27 includes a turbine frame 308 positioned in serialflow arrangement between the first turbine assembly or the high speedturbine 28 and the second turbine assembly or low speed turbine 30. Instill particular embodiments, a bearing assembly 200 is included at theturbine frame 308. Accordingly, the turbine frame 308 may provide astatic mount or support structure at which the bearing assembly 200 ispositioned to support rotation of one or more spools (e.g., low speedspool 36 or high speed spool 34). The turbine frame 308 further includesany appropriate quantity of conduits, manifolds, or passages 309, orother structures for allowing at least a portion of the flow of air 91(e.g., depicted further below as flow of air 193) to the bearingassembly 200. The flow of air to the bearing assembly 200 may providecooling or buffer air at the bearing assembly 200, such as to attenuatevibrations from the spool or generate desired bearing or rotorclearances. In other embodiments, the flow of air 91 is provided to thegear assembly 37 positioned at the fan section 16, to the compressorsection 21, to the turbine section 27, or to the jet exhaust nozzlesection 32.

The engine 10 includes a first conduit 110 extended in fluidcommunication from the compressor section 21 to the turbine section 27.The first conduit 110 is configured to communicate the flow of air 91from the compressor section 21 to a first location 271 at the turbinesection 27. The first conduit 110 forms a flow passage separate from thecore gas flowpath 70. In a particular embodiment, the first conduit 110provides the flow of air 91 from the compressor section 21 to theturbine section 27 while bypassing the combustion section 26.

A first heat exchanger 141 is positioned in thermal communication withthe flow of air 91 through the first conduit 110. The first heatexchanger 141 is configured to receive heat or thermal energy from theflow of air 91 through the first conduit 110. Accordingly, the firstheat exchanger 141 is configured to cool the flow of air 91 through thefirst conduit 110 before the flow of air 91 is provided to the turbinesection 27. The first heat exchanger 141 is configured as anyappropriate heat exchanger for extracting heat or thermal energy fromthe flow of air 91 and receiving or transmitting heat or thermal energyto a heat transfer fluid, depicted schematically via arrows 221.Particular embodiments of the engine 10 may include a fluid system 220configured to flow the heat transfer fluid 221 as a lubricant, a liquidand/or gaseous fuel, a hydraulic fluid, a supercritical fluid, arefrigerant, or an appropriately cooler air or inert gas. The fluidsystem 220 provides the heat transfer fluid 221 into thermalcommunication with the flow of air 91 via the first heat exchanger 141.In a particular embodiment depicted in FIG. 9 (discussed in more detailbelow), the heat transfer fluid 221 is a liquid fuel provided to thecombustion section 26. However, it should be appreciated that the heattransfer fluid 221 may be provided and utilized in any appropriate way,including, but not limited to, as a lubricant for a bearing system, ananti-icing fluid, fuel, or actuation fluid.

Referring still to FIGS. 4-5 , the engine 10 includes a second conduit120 extended from the first conduit 110 downstream of the first heatexchanger 141 (relative to the flow of air 91 from the compressorsection 21 to the turbine section 27). The second conduit 120 isextended in fluid communication to a second location 272 at the turbinesection 27. A flow control device 130 is positioned at the secondconduit 120. The flow control device 130 is configured to selectivelyadjust, alter, modulate, or otherwise change an amount of the flow ofair 91 from the first conduit 110 through the second conduit 120.

In various embodiments, the second conduit 120 includes an inlet portion121 and an outlet portion 122. The inlet portion 121 is fluidly coupledto the first conduit 110 and the flow control device 130. The inletportion 121 extends from the first conduit 110 to provide a portion ofthe flow of air 91, depicted schematically via arrows 192, to the flowcontrol device 130. The outlet portion 122 is fluidly coupled to theflow control device 130 and the second location 272 of the turbinesection 27. The outlet portion 122 extends from the flow control device130 to provide at least a portion of the flow of air 192 to the secondlocation 272 at the turbine section 27. In such a manner, it will beappreciated that for the embodiment depicted, the flow control device130 is positioned between the inlet and outlet portions 121, 122 of thesecond conduit 120.

The flow control device 130 may be a valve or any appropriate device forregulating, directing, controlling, or otherwise modulating an amount offlow of fluid across a passage or flowpath. The flow control device 130may include an actuated valve or an automatic valve driven by anelectric energy source, a pneumatic energy source (e.g., air, orparticularly, at least a portion of the flow of air 91), or a fluidsource (e.g., liquid and/or gaseous fuel, hydraulic fluid, lubricant, orcombinations thereof). The flow control device 130 may include ballvalves, shuttle valves, or other appropriate type of valve or flowregulating device in accordance with the embodiments depicted anddescribed herein. Accordingly, the flow control device 130 is configuredto modulate the amount of flow of fluid through the outlet portion 122of the second conduit 120, such as depicted schematically via arrows 94.

In a particular embodiment, the engine 10 includes a third conduit 123extending from the flow control device 130 to a third location 273 atthe turbine section 27, in fluid communication with both the flowcontrol device 130 and the third location 273. The flow control device130 may therefore be a three-way valve configured to selectively changethe amount of the flow of air 91 from the first conduit 110 through theinlet portion 121 of the second conduit 120 to one or both of the thirdconduit 123 and the outlet portion 122 of the second conduit 120.Accordingly, the flow control device 130 may be configured to modulatean amount of the flow of air 192 through the outlet portion 122 of thesecond conduit 120, such as depicted schematically via arrows 194, andfurthermore modulate and egress of at least a portion of the flow of air192 through the third conduit 123, such as depicted schematically viaarrows 195. The third conduit 123 may form a bypass passage to furtherallow for selective adjustment, control, or modulation of the flows ofair through the flow control device 130. In a particular embodiment, thethird conduit 123 allows for a portion of the air extracted from thefirst conduit 110 to bypass the outlet portion 122 of the second conduit120 and egress to the third location 273 at the turbine section 27. Incertain embodiments, the third location 273 allows for bypassing aclearance control system 275 (described below) and allowing the flow ofair 195 to enter the turbine section 27 at the core gas flowpath 70downstream of the clearance control system 275, or to mix with the flowof air 193 at the turbine frame 308, or to vent to ambient (notdepicted).

Referring still to FIGS. 4-5 , as briefly noted above, the turbinesection 27 includes the clearance control system 275. Exemplaryembodiments of improved clearance control systems are depicted in FIGS.8-16 , including a casing 300, manifold assemblies 316, and thermalcontrol rings 314 such as provided therein. However, it should beappreciated that the clearance control system 275 depicted in FIGS. 4-5may include any appropriate structure or assembly for controlling,adjusting, or otherwise modulating a dimension between a rotor blade tipand a surrounding shroud or wall at the turbine section 27, otherwisereferred to as tip clearance. The clearance control system 275 may be anactive clearance control (ACC) system configured to dynamically controltip clearance. Particularly, the ACC system may be configured todesirably modulate the tip clearance based on an engine operatingcondition via the flow of air 94 received from the second conduit 120and provided to a surrounding shroud at the turbine section 27. Thevolumetric or mass flow rate of the flow of air 94 is regulated ormodulated by the flow control device 130. Modulating the amount of theflow of air 94 to the clearance control system 275 allows the tipclearance to be desirably regulated across various engine operatingconditions and associated changes in temperature at the turbine section27. As temperatures and rotor speeds change at the turbine section 27across various engine operating conditions, the flow control device 130modulates the amount of the flow of air 94 provided to clearance controlsystem 275 to maintain or provide a desired tip clearance. With regardto a landing-takeoff cycle (LTO) of the engine 10 and an aircraft,engine operating conditions include startup, idle, takeoff, climb,cruise, approach, or reverse thrust. However, it should be appreciatedthat other engine operating conditions and cycles may be applicable.

Referring still to FIGS. 4-5 , the second location 272 at the turbinesection 27 is at the clearance control system 275. Accordingly, thesecond conduit 120, or particularly the outlet portion 122 of the secondconduit 120, is fluidly coupled to the turbine section 27 to provide theflow of air 94 to the clearance control system 275 such as describedherein. In a particular embodiment, the clearance control system 275 isoperably coupled to a first turbine assembly or the high speed turbine28 at the turbine section 27. Accordingly, the engine 10 is configuredto receive the flow of air 91 from the compressor section 21 and providethe portion of the flow of air 94 (from the flow of air 91) to theclearance control system 275 at the high speed turbine 28 via the secondconduit 120.

In still particular embodiments, the first conduit 110 is fluidlycoupled to the turbine frame 308 positioned between the first turbineassembly, or the high speed turbine 28, and a second turbine assembly,or low speed turbine 30. The turbine frame 308 may include a pluralityof vanes 310 in circumferential arrangement and positioned between theturbines 28, 30. The first location 271 at the turbine section 27 is atthe turbine frame 308. Accordingly, in such embodiments, the firstconduit 110 is configured to provide at least a portion of the flow ofair 91 to the turbine frame 308 at the first location 271. In aparticular embodiment, schematic arrows 193 depict a portion of the flowof air at the first conduit 110 downstream of a juncture with the secondconduit 120. The flow of air 193 is provided to the turbine frame 308via the first conduit 110. In particular embodiments further depictedand described with regard to FIGS. 8-16 , the flow of air 193 may beprovided to the casing 300 and through the plurality of vanes 310 at theturbine frame 308, such as depicted schematically via arrows 99.

Referring to FIGS. 4-5 , the turbine frame 308 may include or form oneor more passages 309 configured to provide fluid communication of theflow of air 193 to the bearing assembly 200. The flow of air 193 mayprovide a buffer fluid for operation of the bearing assembly 200. Thebuffer fluid may desirably control or attenuate vibrations, or allow orgenerate desired clearances or vibratory responses at the bearingassembly 200 or the rotors to which the bearing assembly is coupled.

Referring now specifically to FIG. 5 , in a particular embodiment, theengine 10 includes a second heat exchanger 142 in thermal communicationwith a flow of air at the bypass airflow passage 48. The second heatexchanger 142 may be configured as a surface heat exchanger configuredto receive heat or thermal energy from the flow of air 194 downstream ofthe flow control device 130 at the second conduit 120. The heat transferfluid at the second heat exchanger 142 is a flow of air through thebypass airflow passage 48 of the engine 10, such as depictedschematically via arrows 177. The second heat exchanger 142 configuredas a surface heat exchanger has a heat exchange surface at the bypassairflow passage 48 and is configured to place the flow of air 194 at thesecond conduit 120 in thermal communication with the flow of bypass air177 at the bypass airflow passage 48. In a particular embodiment, thesecond heat exchanger 142 is positioned at the outlet portion 122 of thesecond conduit 120 and upstream of the second location 272 at theturbine section 27.

Referring back generally to both FIGS. 4-5 , in a particular embodiment,the first conduit 110 includes an inlet manifold 111 configured toreceive the flow of air 91 from a circumferential compressor location211 at the compressor section 21. It should be appreciated that althoughthe embodiments depicted in FIGS. 4-5 depict a single circumferentialcompressor location 211, the inlet manifold may be configured to receivethe flow of air 91 from a plurality of circumferential compressorlocations 211.

Referring now to FIG. 6 , a perspective view of an embodiment of aportion of an engine 10 in accordance with one or more of FIGS. 1through 3 is provided. The embodiment provided in FIG. 6 may beconfigured substantially similarly as described in regard to theembodiments in FIGS. 4-5 . In FIG. 6 , the engine 10 may include aplurality of inlet manifolds 111 evenly-spaced or asymmetrically-spacedalong the circumferential direction C around the compressor section 21.In various embodiments, the plurality of inlet manifolds 111 includestwo (2) or more inlet manifolds. In one embodiment, the plurality ofinlet manifolds 111 includes three (3) inlet manifolds. In anotherembodiment, the plurality of inlet manifolds 111 includes four (4) inletmanifolds and up to 30 inlet manifolds 111.

In FIG. 6 , the first conduit 110 includes a collector 115 configured toreceive the flow of air 91 from the inlet manifold 111. In particularembodiments, the plurality of inlet manifolds 111 is fluidly coupled toa single collector 115 to provide a collected or unified flow of air 91to the first heat exchanger 141. The collector 115 may provide the flowof air 91 to the first heat exchanger 141, such as described herein.

In a still particular embodiment, the first conduit 110 includes anoutlet manifold 112 configured to fluidly communicate the flow of air 91from the first heat exchanger 141 to the turbine section 27 at the firstturbine location 271 at the turbine section 27. The engine 10 mayinclude a plurality of outlet manifolds 112 evenly-spaced orasymmetrically-spaced along the circumferential direction C around theturbine section 27. In various embodiments, the plurality of outletmanifolds 112 includes two (2) or more outlet manifolds. In oneembodiment, the plurality of outlet manifolds 112 includes three (3)outlet manifolds. In another embodiment, the plurality of outletmanifolds 112 includes four (4) outlet manifolds and up to 30 outletmanifolds. In various embodiments, the second conduit 120 is extended influid communication from one or more of the plurality of outletmanifolds 112 of the first conduit 110. The plurality of outletmanifolds 112 may accordingly extend to a plurality of first turbinelocations 271 at different circumferential positions at the turbinesection 27.

It should be appreciated that although the embodiments depicted in FIGS.4-5 depict a single circumferential first turbine location 271, thefirst turbine location 271 may include a plurality of circumferentialfirst turbine locations 271.

Embodiments of the engine 10 provided in FIGS. 4-5 may include the firstconduit 110 as a fixed area flowpath from the compressor section 21 tothe turbine section 27. Stated differently, the first conduit 110 mayinclude various cross-sectional areas or convergent and divergentflowpaths. However, the first conduit 110 and the circumferentialcompressor location 211 may define fixed or non-articulatable flowpathareas. Such fixed area flowpath allows for a constant volumetric or massflow rate of the flow of air 91 from the compressor section 21 throughthe first conduit 110 with respect to a corresponding engine operatingcondition. Stated differently, the fixed area flowpath allows for thefirst conduit 110 to receive a corresponding flow rate of the flow ofair 91 relative to the particular engine operating condition.Accordingly, embodiments of the engine 10 provided herein allow forconstant flows of air 91 in thermal communication with the flow of heattransfer fluid 221 at the first heat exchanger 141. For instance, flowrates of the heat transfer fluid 221, such as a fuel flow rate orlubricant flow rate, may be controlled via a schedule, table, graph, orcurve indicative of the flow rate versus the engine operating condition.In one embodiment, the flow of air 91 at the first conduit 110 maygenerally be fixed as a ratio or proportion of the overall flow of airentering the core engine inlet 20 into the compressor section 21. Inanother embodiment, the flow of air 91 at the first conduit 110 maygenerally be fixed as a ratio or proportion of the flow of air enteringthe high speed compressor 24 from the low speed compressor 22.

The engine 10 may particularly include a variable area flowpath at thesecond conduit 120 via the flow control device 130. Accordingly, theengine 10 may allow a fixed flow of air 193 to the turbine frame 308,such as for the bearing assembly 200, and a variable flow of air 194 tothe clearance control system 275. The flow control device 130 mayadjust, articulate, or otherwise modulate the flow of air 194 to theclearance control system 275 as a function of engine operatingcondition. Modulation of the flow of air 194 via the flow control device130 may be a function of inlet air speed (into the turbomachine 14 viaan inlet 20), or inlet air pressure (e.g., corresponding to altitude ofthe engine 10 during operation or at one or more engine operatingconditions described above), or inlet air temperature, or combinationsthereof. Modulation of the flow of air 194 via the flow control device130 may additionally, or alternatively, be a function tip clearance atthe turbine section 27, or a predetermined schedule corresponding towear or deterioration at the turbine section 27.

Certain embodiments of the engine 10 include particular placements ofthe circumferential compressor location 211 at particular axial stagesor other location at the compressor section 21 corresponding toparticular pressure ranges of the flow of air 91 during operation of theengine 10. In various embodiments, the circumferential compressorlocation 211 from which the flow of air 91 is received from the core gasflowpath 70 corresponds to a compressor location having an airflowtherethrough at a pressure between approximately 20 pounds per squareinch (psi) and approximately psi during an engine operating conditioncorresponding to between approximately 55% and approximately 75% of anoperating envelope. In another embodiment, the circumferentialcompressor location 211 from which the flow of air 91 is received fromthe core gas flowpath 70 may corresponding to a compressor locationhaving an airflow therethrough at a pressure between approximately 30pounds per square inch (psi) and approximately 50 psi during the engineoperating conditions such as described herein.

Accordingly, embodiments of the engine 10 provided herein allow for theclearance control system 275 and the bearing assembly 200 to operate andreceive air from the compressor section 21. In certain embodiments, theengine 10 provided herein allows for the clearance control system 275 toreceive the flow of 91 from the compressor section 21 rather than fromthe bypass airflow passage 48. Furthermore, or alternatively, the engine10 provided herein allows for the flow of air 91 to be received fromupstream, forward, or lower-pressure stages of the compressor section 21in contrast to other compressor bleed systems that may receive highenergy air from downstream, aft, or higher-pressure stages of acompressor section. Certain of these other compressor bleed systems mayfurther mix the higher-energy air with lower-energy (i.e., lowerpressure, lower temperature, or both) corresponding to the bypassairflow passage. Still further, or alternatively, embodiments of theengine 10 provided herein allow for a constant flow of air 91 throughthe first conduit 110 to maintain purge and backflow margin at theturbine frame 308 and bearing assembly 200.

Referring now to FIGS. 7A-7B, a flowchart outlining steps of the method1000 for operating an engine is provided. The steps of the method 1000may be stored as instructions and/or executed as operations byembodiments of the engine 10 and the computing system 1210 providedherein. Accordingly, the method 1000 may be a computer-implementedmethod in which one or more steps is stored as instructions at thememory 1214 at the computing system 1210 and/or executed by one or moreprocessors 1212 at the computing system 1210. The computing system 1210may cause embodiments of the engine such as described herein with regardto FIGS. 1-6 to perform operations such as outlined in the flowchart inFIGS. 7A-7B and described further herein with regard to method 1000.

Referring to the flowchart in FIGS. 7A-7B, and in conjunction with anyone or more embodiments depicted in FIGS. 1-6 , the method 1000 includesat 1010 initiating rotation of one or both of a high speed spool or alow speed spool to, e.g., generate compressed air for combustion withina combustion section of a core engine. In various embodiments, a motiveforce, such as a starter motor or turbine air starter (not shown),initiates rotation of one or both of the high speed spool 34 or the lowspeed spool 36 to generate an initial airflow through the core gasflowpath 70 into the combustion section 26 for mixing with a liquidand/or gaseous fuel before igniting to generate combustion gases.

The method 1000 further includes at 1020 compressing a flow of airthrough the compressor section. During operation of the engine 10, aflow of air 171 is received at the fan section 16. A portion of the flowof air 171 enters the turbomachine 14 through the core engine inlet 20,such as depicted schematically via arrows 172. The flow of air 172 ispressurized across successive rows or stages of compressor blades at thecompressor section 21. Particularly, the low speed compressor 22 mayinclude a low pressure compressor or booster relative to the high speedcompressor 24 including a high pressure compressor. In certainembodiments, a portion of the flow of air 172 compressed by the lowspeed compressor 22 may be bled or re-routed from the core gas flowpath70, such as to control stall, surge, or operability at one or both ofthe compressors 22, 24. The high speed compressor 24 receives the flowof air 172 and further compresses the flow of air, such as depictedschematically via arrows 173 in FIGS. 1-3 . The successive stages ofcompressor blades energize the flow of air 173, such as to increase thepressure and temperature of the flow of air 173 before entering thecombustion section 26, such as depicted via arrows 174.

The method 1000 includes at 1030 extracting a portion of the compressedflow of air from the compressor section, such as described above. Themethod 1000 at 1030 may particularly include extracting the portion ofcompressed flow of air into a first conduit and bypassing a combustionsection, such as provided above with regard to the first conduit 110.The method 1000 includes at 1040 flowing the extracted portion of thecompressed flow of air through the first conduit (e.g., first conduit110) to a turbine section. In a particular embodiment, the first conduitbypasses the combustion section when flowing the extracted portion ofcompressed flow of air to the turbine section. With regard to FIGS. 1-6, a portion of the flow of air at the compressor section 21 is bled orremoved from the core gas flowpath 70 and provided to the first conduit110, such as depicted schematically via arrows 91 in FIGS. 1-5 .Particular embodiments depicted herein may receive the flow of air 91from the compressed flow of air 173, 174 from the high speed compressor24. In still other embodiments, the flow of air 91 may be received fromthe compressed flow of air 172 from the low speed compressor 22.

It should be appreciated that embodiments of the engine 10 providedherein advantageously receive relatively lower-pressure andlower-temperature flows of air from the compressor section 21, and mayfurther avoid structures, complexities, actuatable devices, valves, andassociated weight and efficiency losses related to mixing high-pressureand high-temperature air with low-pressure and low-temperature air fromthe fan bypass airflow passage. It should furthermore be appreciatedthat, while particular operating conditions and operating envelopes areprovided herein, the engine 10 and/or method 1000 provided herein allowsfor performing one or more steps at any engine operating condition,including up to 100% of an overall power output. However, particularadvantages and benefits are provided herein with regard to operation ofthe engine at engine operating conditions defining a majority of anoperating envelope. As such, methods and structures provided hereinallow for improved efficiency and reduced fuel consumption.

In various embodiments, the method 1000 at 1030 includes extracting theportion of the compressed flow of air when the compressed flow of air atthe compressor section is between approximately 20 pounds per squareinch (psi) and approximately 60 psi. In a particular embodiment, themethod 1000 at 1030 includes extracting the portion of the compressedflow of air when the compressed flow of air at the compressor section isbetween approximately 30 psi and approximately 50 psi. In a particularembodiment, the method 1000 includes at 1035 receiving the portion ofthe compressed flow of air from the compressor section, in which theportion of the compressed flow of air is between approximately 20 psiand approximately 60 psi, or between approximately 30 psi andapproximately 50 psi. In a still particular embodiment, the method 1000at 1030 and/or 1035 is performed continuously or constantly relative toa discrete engine operating condition, such as to allow for a fixed flowof air relative to the discrete engine operating condition.

In a still particular embodiment, the method 1000 includes at 1028operating the engine at an engine condition corresponding to betweenapproximately 55% and approximately 75% of an operating envelope, orbetween approximately 60% and approximately 70% of the operatingenvelope, such as described above. In certain embodiments, one or bothsteps of the method 1000 at 1030 and at 1035 is preceded by, orcontemporaneous to, the method 1000 at 1028. In still certainembodiments, the method 1000 includes at 1029 operating the enginebetween approximately 75% and approximately 90% of the overall poweroutput (e.g., rated thrust) of the engine, such as described above. In astill particular embodiment, the method 1000 at 1029 includes operatingthe engine between approximately 80% and approximately 88% of theoverall power output of the engine. In certain embodiments, one or moreranges provided herein may define a discrete engine operating conditionat which the method 1000 at 1030 and/or 1035 is performed continuouslyor constantly. In still particular embodiments, the method 1000 includesperforming the steps at 1028 and 1029 concurrently.

The method 1000 may include at 1050 flowing, via a fluid system, a heattransfer fluid in thermal communication with the extracted portion ofcompressed flow of air, such as described above. In a particularembodiment, the fluid system 220 depicted in FIGS. 4-5 is a liquidand/or gaseous fuel system configured to provide a flow of liquid and/orgaseous fuel to the compressed flow of air 174 to generate combustiongases 175. In such an embodiment, the fuel is the heat transfer fluid221 in thermal communication with the flow of air 91 via the first heatexchanger 141. The flow of fuel receives heat or thermal energy from therelatively hotter flow of air 91, which may advantageously alter certainproperties of the fuel, such as viscosity, density, or other propertythat may desirably affect combustion, fuel-air mixing, swirl, emissionsgeneration, vibrations, or smoke and particulate generation.

In certain embodiments, the method 1000 may further include flowing, viathe fluid system, a plurality of heat transfer fluids in thermalcommunication with the extracted portion of compressed flow of air. Invarious embodiments, the method 1000 includes providing one or moreflows of fuel, lubricant, hydraulic fluid, refrigerant, a supercriticalfluid, or another flow of air at the heat transfer fluid in thermalcommunication with the extracted flow of air.

The method 1000 may further include modulating the flow of the heattransfer fluid to control a temperature of the extracted flow of air(e.g., flow of air 91). Modulating the flow of heat transfer fluid mayinclude adjusting a mass or volumetric flow rate, pressure, ortemperature of the heat transfer fluid provided in thermal communicationwith the extracted flow of air.

As provided above, the flow of liquid and/or gaseous fuel is mixed withthe compressed air from the compressor section and ignited to formcombustion gases 175. The combustion gases 175 flow from the combustionsection 26 to the turbine section 27, and particularly to the high speedturbine 28 and the low speed turbine 30. As the combustion gases 175expand at the turbine section 27, energy is released to drive rotationof the respective turbines 28, 30, which drives their respective spools34, 36, compressors 22, 24, and fan blades 40.

It should be appreciated that the combustion gases 175 release variableamounts of heat at the turbine section 27 based on the engine operatingcondition. Accordingly, heat release and turbine rotor speed may alterthe tip clearance between turbine rotor blade tips and surroundingshrouds, such as further described below. It should be appreciated thatimproved aerodynamic and operating efficiencies are generally achievedby minimizing tip clearances. Accordingly, clearance control systems areutilized to modulate the tip clearance based on engine operatingcondition to improve engine efficiency and performance.

The method 1000 may further include at 1060 selectively flowing aportion of the flow of air through a second conduit (e.g., secondconduit 120) extended from the first conduit (e.g., first conduit 110)downstream of the heat exchanger (e.g., first heat exchanger 141). In aparticular embodiment, the method 1000 includes at 1062 varying ormodulating, via a flow control device (e.g., flow control device 130) atthe second conduit extended from the first conduit, the portion of theflow of air extracted to the second conduit (e.g., second conduit 120)from the first conduit (e.g., first conduit 110) downstream of the heatexchanger (e.g., first heat exchanger 141). In a still particularembodiment, the method 1000 includes at 1063 modulating, via the flowcontrol device, a second portion of the flow of air extracted from thefirst conduit to the third conduit extended from the flow controldevice, such as depicted in FIGS. 4-5 via arrows 195. In a stillparticular embodiment, the method 1000 at 1060 is executedcontemporaneously with the method 1000 at one or more of steps 1028,1030, or 1035. Accordingly, the method 1000 may allow for continuous,constant, or fixed flow of air from the compressor section through thefirst conduit, while modulating or varying the flow of air through thesecond conduit. In particular embodiments, the method 1000 allows forcontinuous, constant, of fixed flow of air from the compressor sectionthrough the first conduit and to the turbine section, or particularlythe bearing assembly, while modulating or varying the flow of airthrough the second conduit to a clearance control system. As such,modulating the flow of air through the second conduit allows for avariable flow of air through to a clearance control system (e.g.,clearance control system 275) independent of whether the operatingcondition of the engine is steady-state (e.g., non-transient ornon-varying) or transient (e.g., changing).

The method 1000 may further include at 1070 selectively varying,altering, or modulating a tip clearance at a clearance control systembased on the flow of air received from the second conduit via step 1060and/or 1062. It should be appreciated that the method 1000 providedherein may further provide for a method for operating a clearancecontrol system and bearing assembly. Such methods may allow for variableflow rate, temperature, pressure, or other physical property of the flowof air through the second conduit to the clearance control system, whileallowing for substantially constant or continuous flows of air throughthe first conduit relative to an engine operating condition.

Although not depicted in FIGS. 7A-7B, the method 1000 may furtherinclude generating a flow of bypass air through a bypass airflowpassage. A portion of the flow of air 171 passes across the plurality offan blades 40 and bypasses the turbomachine 14, such as depicted viaarrows 176 in FIGS. 1-3 . The flow of air 176 that enters the bypassairflow passage 48, depicted schematically via arrows 177, is large involume or mass and cold relative to the flow of air pressurized by thecompressor section 21 within the turbomachine 14. FIG. 5 , which may beapplied to the embodiments of the engine 10 in any of FIGS. 1-3 ,particularly depicts the relatively cold flow of bypass air 177 inthermal communication with the flow of air 194 via the second heatexchanger 142. Accordingly, the method 1000 may further include at 1064thermally communicating, via the second heat exchanger, the flow ofbypass air with the portion of the flow of air extracted to the secondconduit.

The embodiment of the engine 10 depicted and described with regard toFIG. 5 may allow for increased magnitudes of heat transfer from the flowof air 194, such as via the flow of bypass air 177 at the bypass airflowpassage 48. Furthermore, the embodiment depicted in FIG. 5 , whenapplied to an open rotor configuration such as depicted in FIG. 2 , mayovercome challenges associated with removing nacelles and passages,tubes, or conduits that may route through nacelles to provide air forheat exchangers, clearance control systems, and/or bearing assemblies.Accordingly, the method 1000, when applied to an open rotorconfiguration such as described herein, may provide for a method foroperating an open rotor engine, or particularly, a method for operatinga clearance control system for an open rotor engine, or moreparticularly, a method for operating a clearance control system andbearing assembly for an open rotor engine.

Referring now to FIG. 8 , an enlarged cross sectioned view is providedof a turbine section portion of a turbomachine 14 in accordance with oneor more of FIGS. 1-3 , as may incorporate various embodiments of thepresent disclosure. As shown in FIG. 8 , a first turbine assembly isformed by the high speed turbine 28. A first stage 50 of the firstturbine assembly includes a plurality of first turbine rotor blades 58extended within the core gas flowpath 70, and further includes anannular array of stator vanes 54 (only one shown) axially spaced from anannular array of turbine rotor blades 58 (only one shown) at the highspeed turbine 28. In a particular embodiment, the high speed turbine 28further includes a last stage 60 which includes an annular array ofstator vanes 64 (only one shown) axially spaced from an annular array ofturbine rotor blades 68 (only one shown). The turbine rotor blades 58,68 extend radially outwardly from and are coupled to the high speedspool 34 (FIG. 1 , FIG. 2 ). The stator vanes 54, 64 and the turbinerotor blades 58, 68 at least partially define the core gas flowpath 70for routing combustion gases from the combustion section 26 (FIG. 1 ,FIG. 2 ) through the high speed turbine 28.

As further shown in FIG. 8 , the high speed turbine 28 may include oneor more shroud assemblies, each of which forms an annular ring about anannular array of rotor blades. For example, a shroud assembly 72 mayform an annular ring around the annular array of rotor blades 58 of thefirst stage 50 and the annular array of turbine rotor blades 68 of thelast stage 60. In general, the shroud assembly 72 is radially spacedfrom blade tips 76, 78 of each of the rotor blades 58, 68. A radial orclearance gap CL is defined between the blade tips 76, 78 and respectiveinner surfaces of the shroud segments 77. The shroud assembly 72generally reduces leakage from the core gas flowpath 70. The shroudassembly 72 can include a plurality of walls forming thermal controlrings 314 that assist in controlling thermal growth of the shroudthereby controlling the radial deflection or clearance gap CL. Thermalgrowth in the shroud assemblies is actively controlled with theclearance control system 275. The clearance control system 275 is usedto minimize radial blade tip clearance CL between the outer blade tipand the shroud, particularly during cruise operation of the engine, suchas described herein.

Downstream along the core gas flowpath 70, or aft of the high speedturbine 28, is a second turbine assembly formed by the low speed turbine30. As previously described herein, the second turbine assembly isrotatably separate from the first turbine assembly, such as described inregard to the high speed turbine 28 and the low speed turbine 30 abovewith reference to FIG. 1 .

The casing 300 surrounds the high speed turbine 28. The casing 300includes a plurality of vanes 310 extended through the core gas flowpath70 aft of the first turbine assembly formed by the high speed turbine 28and forward of the second turbine assembly formed by the low speedturbine 30. The shroud assembly 72 is coupled to the casing 300 at anouter casing wall 312. The outer casing wall 312 is an annular wallsurrounding the shroud assembly 72 and extended along a circumferentialdirection C relative to the centerline axis 12 (FIGS. 1-3 ). The outercasing wall 312 is extended along an axial direction A forward of therotor blades 58 of the first stage 50 of the high speed turbine 28 (alsoreferred to as the first stage of rotor blades 58) and aft of the rotorblades 68 of the second or last stage 60 of the high speed turbine 28(also referred to as the second stage of rotor blades 68).

The plurality of vanes 310 is extended from the outer casing wall 312.The plurality of vanes 310 is extended into the core gas flowpath 70, Incertain embodiments further described herein, one or more of theplurality of vanes 310 may be hollow or include conduits or passagesallowing for fluid flow within the vane. The outer casing wall 312 ofthe casing 300 is extended along the axial direction A from a downstreamend or trailing edge of the aft-most stage of the rotor blades 68 to atleast an upstream end or leading edge of the plurality of vanes 310,such as depicted at dimension B in FIG. 8 .

It should be appreciated that conventional turbine casings includeseparable or joined flanges, such as bolted flanges or welded flanges,between a high speed turbine casing and a downstream casing, such as aninter-turbine frame, mid-turbine frame, intermediate speed turbinecasing, or low speed turbine casing. Embodiments of the casing 300provided herein, include unitary, integral structures, such as formed byone or more additive manufacturing processes. Embodiments providedherein further form integral, continuous, compliant structures, allowingfor the unitary, integral extension of the casing 300 such as providedherein, or further including one or more features integrally formed tothe casing 300 such as provided herein.

A plurality of walls forming thermal control rings 314 is extended alongthe circumferential direction C and extended outward along a radialdirection R from the outer casing wall 312. In various embodiments, thethermal control rings 314 include forward thermal control rings 3141positioned outward along the radial direction R from the first stage ofrotor blades 58, or particularly from the blade tips 76 of the rotorblades 58, of the high speed turbine 28. In certain embodiments, such asdepicted in FIG. 8 , the forward thermal control rings 3141 arepositioned in alignment along the axial direction A to the first stageof rotor blades 58 (overlapping axial positions). In another particularembodiment, the thermal control rings 314 include aft thermal controlrings 3142 positioned outward along the radial direction R from the laststage 60 of rotor blades 68, or particularly from the blade tips 78 ofthe rotor blades 68, of the high speed turbine 28. In certainembodiments, such as depicted in FIG. 8 , the aft thermal control rings3142 are positioned in alignment along the axial direction A to the laststage 60 of rotor blades 68 of the high speed turbine 28 (overlappingaxial positions).

The forward and aft thermal control rings 3141 and 3142 are provided tomore effectively control blade tip clearance CL (shown in FIG. 8 ) witha minimal amount of time lag and thermal control airflow (cooling orheating depending on operating conditions). The forward and aft thermalcontrol rings 3141 and 3142 are formed with the outer casing wall 312 asan integral, singular, unitary structure of the casing 300. The thermalcontrol rings 314 provide thermal control mass to more effectively movethe shroud segments 77 along the radial direction R to adjust the bladetip clearances CL. Such clearance control may provide for loweroperational specific fuel consumption (SFC).

The integral, unitary structure of the thermal control rings 314 and theouter casing wall 312, with the outer casing wall particularly extendedaft of the second or last stage of the rotor blades 68 of the high speedturbine 28, may allow for improved clearance control, improved thermalcontrol, and improved cooling flow. The structures provided herein allowfor the thermal control rings 314 to be positioned radially outward ofand in axial alignment with each stage of the high speed turbine rotor,such as to improve clearance control at each respective stage. Thestructures provided herein further allow for obviating flanges betweenthe high speed turbine and an intermediate turbine frame between thehigh speed turbine and a downstream low speed turbine (or intermediatespeed turbine, such as described herein).

Embodiments of the integral casing provided herein are generallyproduced by one or more additive manufacturing processes such asdescribed above. Although additive manufacturing may generally beapplied to form various structures or integrate various components, itshould be appreciated that combinations of integrated structuresprovided herein may overcome issues associated with integratingstructures while providing unexpected benefits. In one instance,axially-extended casings may generally be susceptible to thermaldistortion that may ovalize the core flowpath, which may adverselyaffect rotor operation as the rotors may rub within a non-concentricflowpath. As such, simple integration of relatively hot casingssurrounding the high speed turbine with the relatively cooler casingsurrounding downstream vanes proximate to the low speed turbine mayadversely affect overall engine operation. In another instance, suchlarge, axially-extended masses may require additional cooling flow,which results in increased fuel consumption and overall losses in engineperformance.

Embodiments of the engine provided herein overcome such issues at leastin part by the positioning of the thermal control rings in axialalignment and radially outward of the respective stages of the highspeed turbine blades. Removing flanges between a casing surrounding thehigh speed turbine rotors and a vane casing or frame downstream of thehigh speed turbine allows for the thermal control rings to beadvantageously positioned as disclosed herein.

Other embodiments of the engine provided herein overcome such issues atleast in part by improved cooling flow structures, passages, andconduits. In various embodiments, a manifold assembly 316 surrounds thethermal control rings 314 along the circumferential direction C and theaxial direction A. The manifold assembly 316 is configured to provide aflow of fluid, such as the flow of air 192 from the compressor section21 such as depicted and described in regard to FIGS. 4-5 , to thethermal control rings 314.

Referring still to FIG. 8 and now also to FIGS. 9-11 , and FIG. 14 ,further exemplary embodiments are provided. The embodiment depicted inFIG. 8 , FIG. 9 , and FIG. 14 may be configured similarly as oneanother, such as further described below. FIGS. 9-11 provide views offlows of fluid and openings at various cross-sections of an embodimentof the engine 10 at different circumferential positions of the engine10. Each of the embodiments may be formed via one or more manufacturingmethods known in the art. In FIG. 14 , the embodiment provided mayinclude double-wall structures that may be formed via an additivemanufacturing process. Various embodiments provided herein may be formedas integral, unitary structures, such as via an additive manufacturingprocess or other appropriate manufacturing process.

Referring to the various embodiments depicted in FIGS. 8-11 and FIG. 14, the manifold assembly 316 is extended along the axial direction Aforward and aft of the plurality of axially-spaced stages of theplurality of walls forming the thermal control rings 314. In aparticular embodiment, such as depicted in FIG. 14 , the manifoldassembly 316 is extended aft along the axial direction A of theplurality of vanes 310. In various embodiments, such as in the exemplaryembodiment of FIG. 8 , the manifold assembly 316, the outer casing wall312, and the plurality of walls forming the thermal control rings 314 ofthe casing 300 is a single, integral, unitary structure, such asdescribed herein. In particular embodiments, such as in the exemplaryembodiment of FIG. 8 , the manifold assembly 316 includes a plurality ofconcentric walls integrally formed and surrounding the outer casing wall312. In certain embodiments, the manifold assembly 316 includes an innermanifold 1316 radially inward of and concentric to an outer manifold2316. In still certain embodiments, the inner manifold 1316 is a doublewall structure concentric to the outer manifold 2316.

Referring particularly to FIGS. 9-10 , certain embodiments of the casing300 include a corrugated feature 399. The corrugated feature 399includes a shape defining ridges or grooves configured to mitigateformation of thermal expansion stresses at the casing 300. In certainembodiments, the corrugated feature 399 is formed at the manifoldassembly 316. In a still particular embodiment, the corrugated feature399 may be formed at an inner manifold 1316 or an outer manifold 2316.The corrugated feature 399 may allow for the unitary, integral formationof the manifold assembly 316 with the outer casing wall 312, such asdescribed in various embodiments herein.

Referring now briefly to FIG. 15 , the manifold assembly 316 includes aplurality of openings 318 surrounding the plurality of walls forming thethermal control rings 314 at the casing 300. The plurality of openings318 allow for the flow of fluid, depicted schematically via arrows 91,to come into thermal communication with the thermal control rings 314for desired heat transfer effect. In various embodiments, the pluralityof openings 318 include an inlet opening 3181 configured to allow theflow of air 91 into a first cavity 1321 in thermal communication withthe thermal control rings 314, as described further below. The pluralityof openings 318 may further include an outlet opening 3182 configured toallow at least a portion of the flow of air 91, depicted schematicallyvia flow of air 92, to egress the first cavity 1321 and enter an innerwall conduit 1326 such as described further below.

An inlet opening wall 381 is extended between an outer portion 346 andan inner portion 347 of the double wall structure formed by the innermanifold 1316. The inlet opening wall 381 forms an inlet openingflowpath 382 that extends through the double wall structure fluidlyseparated from the inner wall conduit 1326. The inlet opening 3181 andthe inlet opening wall 381 allow for the flow of air 91 to pass from aconduit 1324 surrounding the inner manifold 1316 to enter a plenum 383formed between adjacent thermal control rings 314. Particularly, theinlet opening wall 381 extends between the outer portion 346 and innerportion 347 of the inner manifold 1316. The inlet opening flowpath 382formed by the inlet opening wall 381 allows the flow of air 91 to enterthe plenum 383 while being fluidly segregated from the flow of air 92through the inner wall conduit 1326.

Referring back particularly to FIGS. 9-10 , as discussed, the manifoldassembly 316 includes the inner manifold 1316 surrounding the thermalcontrol rings 314 along the circumferential direction C and the axialdirection A. The manifold assembly 316 depicted further includes theouter manifold 2316 surrounding the inner manifold 1316, as discussedabove. A passage wall 1318 is extended to the outer manifold 2316 fromthe inner manifold 1316 to form a passage 1320 within the passage wall1318.

In certain embodiments, such as depicted in FIG. 8 , the outer manifold2316 of the manifold assembly 316 is extended along the axial directionA at or aft the plurality of vanes 310. The outer manifold 2316 isfurther connected to the outer casing wall 312 at or aft of theplurality of vanes 310. In still certain embodiments, such as depictedin FIGS. 9-11 , the inner manifold 1316 is extended to a locationforward along the axial direction A of the plurality of vanes 310(terminating forward of the plurality of vanes 310). The inner manifold1316 is also extended to a location aft along the axial direction A ofthe plurality of walls forming the thermal control rings 314. As such,the inner manifold 1316 is connected to the outer casing wall 312forward of the plurality of vanes 310 and aft of the thermal controlrings 314.

The first cavity 1321 discussed above with reference to FIG. 15 (alsodepicted in FIGS. 9-11 ) is formed between the inner manifold 1316 andthe outer casing wall 312. The thermal control rings 314 are surroundedby the inner manifold 1316 at a location within the first cavity 1321between the inner manifold 1316 and the outer casing wall 312. Thepassage 1320 allows for fluid communication with the first cavity 1321between the inner manifold 1316 and the outer casing wall 312. Thepassage 1320 further allows for the flow of air 91 to enter into thermalcommunication with the thermal control rings 314.

In various embodiments, the conduit 1324 briefly mentioned above isformed between the outer manifold 2316 and the inner manifold 1316. Theconduit 1324 is in fluid communication with the first cavity 1321 and isfluidly separated from passage 1320 by the passage wall 1318. Inparticular embodiments, the passage wall 1318 is extended from the outermanifold 2316 to the inner manifold 1316 through the conduit 1324.

Referring particularly to FIGS. 9-11 , and further in regard to FIG. 14, the conduit 1324 is further extended in fluid communication throughone or more of the plurality of vanes 310. FIG. 10 and FIG. 14particularly depict the flow of air 91 entering into thermalcommunication and fluid communication with the thermal control rings 314in the first cavity 1321. FIG. 10 particularly depicts the flow of air91 entering into thermal communication and fluid communication with thethermal control rings 314 in the first cavity 1321. In variousembodiments, the first cavity 1321 is formed to direct the flow of fluidto thermal contact portions of the thermal control rings directly, suchas in a perpendicular direction. FIG. 11 and FIG. 14 particularly depictthe flow of air 92 egressing from the first cavity 1321 through theconduit 1324 and then in serial flow through one or more of theplurality of vanes 310 (as airflow 99, discussed below). In certainembodiments, the thermal control rings 314 are formed with the outercasing wall 312 to desirably improve clearance control. In oneembodiment, such as depicted in FIG. 13B, the thermal control ring 314includes outer surfaces extended as a ridge, groove, or at acute orzig-zagging angles (see more detailed description below).

Referring briefly particularly to FIG. 14 , and further depicted in thedetailed perspective view in FIG. 15 , in certain embodiments, the innermanifold 1316 is a double wall structure forming the inner wall conduit1326 between the double wall structure of the inner manifold 1316. Theinner wall conduit 1326 may extend in fluid communication to a secondcavity 1322 formed between the outer casing wall 312 and an outer wall170 of the core gas flowpath 70. In such embodiments, the unitary,integral casing 300, or furthermore integral to embodiments of themanifold assembly 316, allow for separate flows into the plurality ofvanes 310. Particularly, the flow of air 91 enters the conduit 1324 froma compressor section such as depicted and described with regard to FIGS.1-6 . A portion of the flow of air 91, depicted via arrows 92, flowsinto the first cavity 1321 and then into the inner wall conduit 1326formed at the double wall structure. The flow of air 92 then flows intoone or more of the plurality of vanes 310. Furthermore, another portionof the flow of air 91, depicted via arrows 99, remains in the conduit1324 and flows into one or more of the plurality of vanes 310. Incertain embodiments, the flows 92, 99 are isolated or fluidly separatedfrom one another until mixing at the plurality of vanes 310. In otherembodiments, the flows 92, 99 remain fluidly separated and are providedto separate respective vanes 310, or separate conduits within each vane310. Embodiments of the casing 300 and the manifold assembly 316 allowfor improved thermal efficiency and improved overall engine efficiency,such as via providing secondary uses of the flow of fluid after thermalcommunication with the thermal control rings 314, rather than outputtingthe flows to atmosphere or to an under-cowl area of the engine.

In certain embodiments, the outer wall 170 of the core gas flowpath 70forms the outer shroud segment 77 of the shroud assembly 72. The outershroud segment 77 is exposed to the core gas flowpath 70, and mayinclude thermal barrier coatings or materials configured to withstandheat from the combustion gases. The outer shroud segment 77 may furtherbe configured to at least partially rub with one or more stages ofblades at the core gas flowpath 70.

Referring still to FIG. 14 , and further depicted in FIG. 15 , FIG. 16providing a side view of the casing 300 of FIG. 15 , and FIG. 17providing a close-up view of Section A in FIG. 16 , the inner manifold1316 includes a plenum wall 1319 extended from the inner manifold 1316and surrounding the thermal control ring 314. In certain embodiments,the plenum wall 1319 is extended radially inward from the inner portion347 of the inner manifold 1316. The plenum wall 1319 may be formed as anintegral, unitary, or monolithic structure with the inner manifold 1316including the outer portion 346 and the inner portion 347. The firstcavity 1321 is formed between an outer surface 1325 of the thermalcontrol ring 314 and the plenum wall 1319.

Referring particularly to FIGS. 16 and 17 , the thermal control ring 314includes a wall or body 332 extended outward, such as outward along theradial direction R, from the outer casing wall 312. In variousembodiments, such as described above with regard to the plurality ofthermal control rings 314, the body 332 is extended substantiallyannularly along the circumferential direction C (FIGS. 1-3 ).

Referring more particularly to FIG. 17 , the body 332 forms an internalflowpath 330 to allow a flow of fluid through the thermal control ring314. The flow of fluid through the body 332 allows for a temperature orthermal gradient at the thermal control ring 314 to be desirablycontrolled, altered, or modulated by changes in temperature or flow rateof the flow of fluid through the flowpath 330 at the body 332. The flowof fluid through the body 332 may furthermore allow for one or morestructures attached or integrally formed to the thermal control ring314, such as the outer casing wall 312 or the shroud assembly 72, tomove based at least in part on thermal changes provided by the flow offluid, such as to desirably control the clearance gap CL (FIG. 8 )between the rotor blades 58, 68 and the shroud assembly 72.

Referring still to FIG. 17 , the exemplary casing 300 depicted furtherincludes a plurality of pins 334 extended along a radial direction R ofthe engine 10 incorporating the casing 300 from the outer casing wall312 to the body 332. Referring briefly also to FIG. 18 , a top-down viewof the plurality of pins 334 depicts each pin 334 is depicted. As shownin FIGS. 17 and 18 , each pin 334 spaced apart from one another along anaxial direction A of the engine 10 incorporating the casing 300 andalong a circumferential direction C of the engine 10 incorporating thecasing 300 (FIG. 18 ). In such a manner, adjacent pins 334 define aspace 336 therebetween.

Referring back particularly to FIG. 17 , the flowpath 330 extendedradially through the body 332 is further extended in fluid communicationto the gap or space 336 provided between the plurality of pins 334. Thethermal control ring 314 may form the flowpath 330 as a plurality ofdiscrete, round or slotted flowpaths in adjacent arrangement along thecircumferential direction C. In other embodiments, the thermal controlring 314 forms the flowpath 330 as a plurality of arcuate sectionsextended at least partially along the circumferential direction C. Theflow of air, depicted schematically via arrows 91 is received andprovided in fluid communication with the thermal control rings 314 inaccordance with any one or more embodiments depicted and described abovewith regard to FIGS. 1-15 .

During operation, the flow of air 91 passes through the spaces 336 andacross the plurality of pins 334 to enter into the flowpath 330 withinthe body 332. During operation, the flow of air 91 progresses radiallythrough the body 332 and egresses the body 332 through an outlet opening338 at the flowpath 330. The outlet opening 338 is formed by the body332 distal to the spaces 336 to allow for fluid communication from theflowpath 330 to the inner wall conduit 1326 formed within the doublewall structure of the inner manifold 1316. The flow of fluid egressedfrom the thermal control ring 314, depicted schematically via arrows 92,may flow through the inner wall conduit 1326 in accordance with any oneor more embodiments depicted and described with regard to FIGS. 1-15 .

Referring still to FIG. 17 , in various embodiments, a seal 1323 ispositioned to contact the outer surface 1325 of the thermal control ring314 and the plenum wall 1319. Additionally, or alternatively, the seal1323 may be formed or positioned in contact with the inner portion 347of the inner manifold 1316 and the outer surface 1325 of the body 332 ofthe thermal control ring 314. The seal 1323 inhibits a flow of fluidthrough the first cavity 1321. In a particular embodiment, the seal 1323may form a structural member configured to provide structural support tothe inner manifold 1316 and/or the thermal control ring 314. The seal1323 may further support the body 332 relative to the plurality of pins334. In certain embodiments, the seal 1323 is a braze, weld, or othermember attaching the plenum wall 1319 to the thermal control ring 314 atthe first cavity 1321. It should be appreciated that the seal 1323 andthe plenum wall 1319 may each extend substantially co-directional withthe thermal control ring 314 as either a monolithic annular component oras a plurality of arcuate sections positioned in circumferentialarrangement.

In particular embodiments, the outer casing wall 312, the plurality ofpins 334, and the body 332 of the thermal control rings 314 are aunitary, integral structure, such as may be formed by an additivemanufacturing process, or other appropriate manufacturing process. Instill particular embodiments, the inner portion 347, the outer portion346, and the plenum wall 1319 are together formed as a unitary, integralstructure of the inner manifold 1316. In certain embodiments, thethermal control rings 314 and outer casing wall 312 are a unitarystructure separate from the inner manifold 1316. In still certainembodiments, the unitary structures are formed from an additivemanufacturing process.

Referring now to FIG. 19 , an exemplary embodiment is provided depictingan operation of the engine 10. The embodiment provided in FIG. 19 isconfigured substantially similarly to the embodiment depicted anddescribed with regard to FIG. 16 . Operation of the system provided heremay be based substantially as described with regard to embodiments ofthe engine 10 as depicted and described with regard to FIGS. 1-6 andFIGS. 7A-7B. In FIG. 19 , the flow of air 91 is received at the secondlocation 272, such as an opening provided through the outer manifold2316. The flow of air 91 is received into the conduit 1324 formedbetween the outer manifold 2316 and the inner manifold 1316. The flow ofair 91 is routed into the plenum 383 via the inlet opening 1381 formedthrough the inner manifold 1316. The flow of air 91 is routed across theplurality of pins 334 and through the flowpath 330 (see FIG. 17 ) intothe inner wall conduit 1326 (see FIG. 17 ).

In one embodiment, such as depicted in FIG. 19 , the flow of air 92 mayegress from the inner wall conduit 1326 to outside of the casing 300 orengine 10, such as depicted via arrows 93 through opening 1380. The flowof air 93 may egress heat or thermal energy from the thermal controlrings 314 to an atmospheric condition, or to an under-casing orunder-cowl area.

Referring now to FIG. 20 , a perspective view of a portion of the engine10 is provided. The embodiment provided in FIG. 20 is configuredsubstantially similarly to the embodiment described with regard to FIGS.16-19 . In particular, FIG. 20 depicts a plurality of discrete flowpaths330 extended in adjacent circumferential arrangement through the thermalcontrol rings 314. A plurality of outlet openings 3182 is formed throughthe inner portion 347 of the inner manifold 1316 corresponding to theplurality of flowpaths 330 and outlet openings 338 at the thermalcontrol rings 314. The engine 10 may accordingly form a plurality offlowpaths 330 and outlet openings 338 at the thermal control rings 314in adjacent arrangement along the circumferential direction Ccorresponding to the plurality of outlet openings 3182 formed throughthe inner portion 347 of the inner manifold 1316. Such arrangement mayallow for the flow of air 92 to egress from within the thermal controlring 314 into the inner wall conduit 1326.

Referring now to FIG. 21 , a side cross-sectional view of the embodimentprovided in FIG. 20 is provided. The embodiment in FIG. 21 furtherdepicts the inner wall conduit 1326 in fluid communication with thesecond cavity 1322 positioned at the turbine frame 308. An opening 3112is formed through the turbine frame 308 to allow the flow of air 92 toegress into thermal communication with the turbine frame 308.

Referring briefly now back to FIG. 12 and FIGS. 13A-13D, additionalaspects of the present disclosure are described. FIG. 12 provides apartial circumferential view of an embodiment of the manifold assembly316. FIGS. 13A-13D furthermore provide sectional views of the embodimentdepicted in FIG. 12 (labels for each of FIGS. 13A-13D are indicated inFIG. 12 ). As previously described, various embodiments of the manifoldassembly 316 are formed via one or more additive manufacturingprocesses. Referring particularly to FIG. 12 and the close-up view ofFIG. 13C, in various embodiments, a member 3316 is extended to the innermanifold 1316 and the outer manifold 2316. The member 3316 is extendedat an acute angle (e.g., a V-, Z-, or other angled cross-section) fromthe inner manifold 1316 to the outer manifold 2316. In variousembodiments, the member 3316 is extended along a first direction,depicted schematically via arrows 95, and a second direction opposite ofthe first direction, depicted schematically via arrows 96.

Embodiments of the improved turbine casing 300, turbine section 27, andengine 10 provided herein allow for improved clearance control, coolingfluid distribution, reduced weight, and improved engine efficiency.Embodiments of the engine 10, the casing 300, and manifold assembly 316provided herein include integral, unitary structures, such as the casingextended over the stages of the high speed turbine, or further includingthe inter-turbine frame, or further including all or part of themanifold, such as may be formed by additive manufacturing processes thatwould not have heretofore been possible or practicable. Embodimentsdepicted and described herein allow for improved and advantageouspositioning of thermal control rings 314, flowpaths 330 therethrough,and the plurality of pins 334, for improved clearance control response,improved formation and positioning of openings, passages, and conduitsto allow for more efficient heat transfer fluid utilization andmovement, and reduced weight, such as via obviating flanges andsub-assemblies into integral components. Particular combinations ofthese features allow for improved heat transfer properties and reducedthermal gradients. Improved heat transfer properties particularlyinclude lowering a heat transfer coefficient at certain features, suchas the plurality of walls, body, pins, and/or flowpaths forming thethermal control rings 314, in contrast to known clearance controlsystems. Such improvements may mitigate or eliminate undesired orexcessive deformation, ovalization, bowing, or other changes in geometryof the casing 300 that may adversely affect deflections or result inundesired contact to the turbine rotor blades 58 at the high speedturbine 28.

Embodiments of the engine 10 and the casing 300 provided herein includean integral, unitary casing for the high speed turbine 28 together witha turbine center frame or mid-turbine frame 308, formed by the outercasing wall 312 and the plurality of vanes 310 and positioned downstreamalong the core gas flowpath 70 of the high speed turbine 28 and upstreamalong the core gas flowpath 70 of a low- or intermediate-pressureturbine, such as depicted at turbine 30. Embodiments provided hereinfurther include e.g., an integral, unitary clearance control manifoldconfigured to provide heat transfer fluid to thermal control rings. Theintegral, unitary structures may further allow for improved positioningof the thermal control rings relative to the turbine rotors, such as toprovide improved clearance control across the turbine rotor assembly.

It should be appreciated that the conduits 110, 120, 123, flow controldevices 130, or heat exchangers 141, 142 depicted and described withregard to FIGS. 1-6 may be provided to the casing 300, manifold assembly316, and other structures depicted and described with regard to FIGS.8-21 . However, various embodiments of the engine 10 provided herein mayinclude one or more of the conduits 110, 120, 123, flow control devices130, or heat exchangers 141, 142 providing flows of air to anyappropriate clearance control system, turbine section, or bearingassembly. Such structures, when combined with any appropriate clearancecontrol system, turbine section, or bearing assembly, may provide one ormore advantages and benefits described herein. Alternatively, variousembodiments of the engine 10 provided herein may include one or more ofthe casings 300 or the manifold assemblies 316 receiving flows of airfrom any appropriate conduits, passageways, flowpaths, tubes, or otherstructures. Such structures, when combined with any appropriate conduitor heat exchanger, may provide one or more advantages and benefitsdescribed herein. Benefits and advantages described with regard toeither the conduits, flow control devices, heat exchangers, casings, ormanifolds, when combined together, may compound such benefits andadvantages described herein.

Embodiments of the conduits 110, 120, 123 and heat exchangers 141, 142provided herein may be formed, at least in part, by one or more additivemanufacturing processes such as described herein. For instance, thefirst heat exchanger 141 may be integrally formed with the first conduit110, or the second heat exchanger 142 may be integrally formed with thesecond conduit 120, or portions thereof. In another instance, all orpart of the first conduit 110, including one or more inlet manifolds111, outlet manifolds 112, or collectors 115 may be integrally formed asa single, unitary component. In still another instance, all or part ofthe second conduit 120, including one or more inlet portions 121 oroutlet portions 122 may be formed as a single, unitary component. Stillfurther, certain combinations of portions of the first conduit 110,second conduit 120, and third conduit 123 may be formed integrally toone another. For instance, the outlet manifold 112 may be formed as asingle, unitary component with the inlet portion 121. In anotherinstance, casings surrounding the compressor section 21 may be formedintegrally with the inlet manifold 111. The collector 115 may be formedintegrally with the first heat exchanger 141. The second heat exchanger142 may be formed integrally with the outlet portion 122.

This written description uses examples to disclose the preferredembodiments, including the best mode, and also to enable any personskilled in the art to practice the disclosure, including making andusing any devices or systems and performing any incorporated methods.The patentable scope of the disclosure is defined by the claims, and mayinclude other examples that occur to those skilled in the art. Suchother examples are intended to be within the scope of the claims if theyinclude structural elements that do not differ from the literal languageof the claims, or if they include equivalent structural elements withinsubstantial differences from the literal languages of the claims.

Further aspects of the disclosure are provided by the subject matter ofthe following clauses:

A method of operating a gas turbine engine having a compressor sectionand a turbine section in serial flow arrangement, a first conduit influid communication with the compressor section and the turbine section,a first heat exchanger positioned in thermal communication with a flowof air through the first conduit, a second conduit in fluidcommunication with the first conduit at a location downstream of thefirst heat exchanger and in fluid communication with a second locationat the turbine section, and a flow control device positioned in flowcommunication with the second conduit, the method comprising: extractingthe flow of air from the compressor section into the first conduit;flowing the extracted flow of air through the first conduit to the afirst location at the turbine section, wherein the second conduit is influid communication with the turbine section at a second location;flowing a heat transfer fluid to the first heat exchanger, the heattransfer fluid in thermal communication with the extracted flow of airthrough the first conduit via the first heat exchanger; and modulating,via the flow control device, a portion of the flow of air extracted fromthe first conduit to the second conduit downstream of the first heatexchanger.

The method of one or more of these clauses, further comprising:operating the engine between approximately 75% and approximately 90% ofan overall power output of the engine.

The method of one or more of these clauses, wherein extracting the flowof air from the compressor section comprises extracting the flow of airfrom the compressor section at a compressor location having an airflowpressure between approximately 20 pounds per square inch andapproximately 60 pounds per square inch while operating the enginebetween approximately 75% and approximately 90% of the overall poweroutput.

The method of one or more of these clauses, further comprising:operating the engine at an operating condition corresponding to betweenapproximately 55% and approximately 75% of an operating envelope; andwherein extracting the flow of air from the compressor section comprisesextracting the flow of air from the compressor section at a compressorlocation having an airflow pressure between approximately 20 pounds persquare inch and approximately 60 pounds per square inch during theoperating condition corresponding to between approximately 55% andapproximately 75% of the operating envelope.

The method of one or more of these clauses, wherein the compressorsection comprises a low speed compressor and a high speed compressor,and wherein the turbine section comprises a low speed turbine, a highspeed turbine, and a turbine frame positioned between the low speedturbine and the high speed turbine, and wherein the first location atthe turbine section is at the turbine frame, and wherein the firstconduit extends in fluid communication from the high speed compressor tothe turbine frame.

The method of one or more of these clauses, wherein the first conduit isconfigured as a fixed area flowpath from the compressor section to theturbine section, and wherein the flow control device defines a variablearea flowpath at the second conduit, and wherein flowing the extractedflow of air through the first conduit to the first location at theturbine section is a continuous flow through an operating condition ofthe engine, and wherein modulating the portion of the flow of airextracted from the first conduit to the second conduit comprisesproviding a variable flow to the second location at the turbine section.

The method of one or more of these clauses, wherein the gas turbineengine further comprises: a clearance control system positioned at thesecond location at the turbine section; and wherein the method furthercomprises: selectively altering a tip clearance at the clearance controlsystem based on the portion of the flow of air extracted from the firstconduit to the second conduit.

The method of one or more of these clauses, wherein the gas turbineengine comprises: a third conduit extended from the flow control deviceand in fluid communication with a third location at the turbine section;wherein the method further comprises: modulating, via the flow controldevice, a second portion of the flow of air extracted from the firstconduit to the third conduit extended from the flow control device.

The method of one or more of these clauses, wherein the gas turbineengine further comprises: a fan section, wherein a bypass airflowpassage is formed downstream of the fan section and around an outercasing surrounding the compressor section and the turbine section; and asecond heat exchanger positioned at the second conduit downstream of theflow control device and upstream of the second location at the turbinesection, wherein the second heat exchanger allows for thermalcommunication of a flow of bypass air from the bypass airflow passagewith the portion of the flow of air extracted to the second conduit;wherein the method further comprises: thermally communicating, via thesecond heat exchanger, the flow of bypass air with the portion of theflow of air extracted to the second conduit.

A gas turbine engine, the gas turbine engine comprising: a compressorsection and a turbine section in serial flow arrangement; a firstconduit extended from the compressor section to the turbine section, thefirst conduit in fluid communication with the compressor section and theturbine section to communicate a flow of air from the compressor sectionto a first location at the turbine section; a first heat exchangerpositioned in thermal communication with the flow of air through thefirst conduit; a second conduit in fluid communication with the firstconduit at a location downstream of the first heat exchanger and influid communication with a second location at the turbine section; and aflow control device positioned in flow communication with the secondconduit for selectively changing an amount of the flow of air from thefirst conduit through the second conduit.

The gas turbine engine of one or more of these clauses, wherein thesecond conduit comprises an inlet portion and an outlet portion, whereinthe inlet portion is fluidly coupled to the first conduit and the flowcontrol device, and wherein the outlet portion is fluidly coupled to theflow control device and the second location of the turbine section.

The gas turbine engine of one or more of these clauses, the gas turbineengine comprising: a second heat exchanger in thermal communication witha flow of air through the outlet portion of the second conduit andupstream of the second location at the turbine section.

The gas turbine engine of one or more of these clauses, the gas turbineengine comprising: a fan section, wherein a bypass airflow passage isformed downstream of the fan section and around an outer casingsurrounding the compressor section and the turbine section, wherein thefan section provides a flow of bypass air to the bypass airflow passageduring operation of the gas turbine engine, and wherein the second heatexchanger provides the flow of bypass air at the bypass airflow passageinto thermal communication with the flow of air at the outlet portion ofthe second conduit.

The gas turbine engine of one or more of these clauses, the gas turbineengine comprising: a third conduit extended in fluid communication fromthe flow control device to a third location at the turbine section.

The gas turbine engine of one or more of these clauses, wherein the flowcontrol device is a three-way valve configured to selectively change theamount of the flow of air from the first conduit through an inletportion of the second conduit, and wherein the flow control device isconfigured to egress at least a portion of the flow of air to the thirdconduit, an outlet portion of the second conduit, or both.

The gas turbine engine of one or more of these clauses, the turbinesection comprising: a first turbine assembly surrounded by an outer wallforming a gas flowpath, and wherein the first turbine assembly comprisesan outer casing wall surrounding the outer wall, wherein the outer walland the outer casing wall together form a cavity, and wherein the thirdlocation at the turbine section is at the cavity.

The gas turbine engine of one or more of these clauses, the turbinesection comprising: a first turbine assembly; a clearance controlsystem, wherein the second location at the turbine section is at theclearance control system and wherein the clearance control system isoperably coupled to the first turbine assembly; and a turbine framepositioned in serial flow arrangement downstream of the first turbineassembly, and wherein the first location at the turbine section is atthe turbine frame.

The gas turbine engine of one or more of these clauses, the gas turbineengine comprising: a fluid system configured to provide a flow of heattransfer fluid in thermal communication with the flow of air via thefirst heat exchanger.

The gas turbine engine of one or more of these clauses, wherein thefirst conduit comprises: a plurality of inlet manifolds configured toreceive the flow of air from a plurality of circumferential compressorlocations at the compressor section; and a collector configured toreceive the flow of air from the plurality of inlet manifolds, andwherein the collector is fluidly coupled to the first heat exchanger toprovide the flow of air in thermal communication to the first heatexchanger.

An airflow delivery system for a gas turbine engine, the gas turbineengine comprising a compressor section and a turbine section in serialflow arrangement, the airflow delivery system comprising: a firstconduit configured to extend from the compressor section to the turbinesection in flow communication with the compressor section and theturbine section to communicate a flow of air from the compressor sectionto a first location at the turbine section when installed in the gasturbine engine; a first heat exchanger positioned to be in thermalcommunication with the flow of air at the first conduit; a secondconduit extending from the first conduit downstream of the first heatexchanger, wherein the second conduit is configured to extend in fluidcommunication to a second location at the turbine section; and a flowcontrol device positioned at the second conduit, wherein the flowcontrol device is configured to selectively change an amount of the flowof air from the first conduit through the second conduit.

What is claimed is:
 1. A gas turbine engine, the gas turbine enginecomprising: a compressor section and a turbine section in serial flowarrangement; a first conduit extended from the compressor section to theturbine section, the first conduit in fluid communication with thecompressor section and the turbine section to communicate a flow of airfrom the compressor section to a first location at the turbine section;a first heat exchanger positioned in thermal communication with the flowof air through the first conduit; a second conduit in fluidcommunication with the first conduit at a location downstream of thefirst heat exchanger and in fluid communication with a second locationat the turbine section, the second conduit configured to receive aportion of the flow of air from the first conduit, and wherein thesecond conduit is configured to deliver only the portion of the flow ofair to the second location at the turbine section; and a flow controldevice positioned in flow communication with the second conduitconfigured to modulate an amount of the flow of air from the firstconduit through the second conduit.
 2. The gas turbine engine of claim1, wherein the second conduit comprises an inlet portion and an outletportion, wherein the inlet portion is fluidly coupled to the firstconduit and the flow control device, and wherein the outlet portion isfluidly coupled to the flow control device and the second location ofthe turbine section.
 3. The gas turbine engine of claim 2, the gasturbine engine comprising: a second heat exchanger in thermalcommunication with a flow of air through the outlet portion of thesecond conduit and upstream of the second location at the turbinesection.
 4. The gas turbine engine of claim 3, the gas turbine enginecomprising: a fan section, wherein a bypass airflow passage is formeddownstream of the fan section and around an outer casing surrounding thecompressor section and the turbine section, wherein the fan sectionprovides a flow of bypass air to the bypass airflow passage duringoperation of the gas turbine engine, and wherein the second heatexchanger provides the flow of bypass air at the bypass airflow passageinto thermal communication with the flow of air at the outlet portion ofthe second conduit.
 5. The gas turbine engine of claim 1, the gasturbine engine comprising: a third conduit extended in fluidcommunication from the flow control device to a third location at theturbine section.
 6. The gas turbine engine of claim 5, wherein the flowcontrol device is a three-way valve configured to selectively change theamount of the flow of air from the first conduit through an inletportion of the second conduit, and wherein the flow control device isconfigured to egress at least a portion of the flow of air to the thirdconduit, an outlet portion of the second conduit, or both.
 7. The gasturbine engine of claim 5, the turbine section comprising: a firstturbine assembly surrounded by an outer wall forming a gas flowpath, andwherein the first turbine assembly comprises an outer casing wallsurrounding the outer wall, wherein the outer wall and the outer casingwall together form a cavity, and wherein the third location at theturbine section is at the cavity.
 8. The gas turbine engine of claim 1,the turbine section comprising: a first turbine assembly; a clearancecontrol system, wherein the second location at the turbine section is atthe clearance control system and wherein the clearance control system isoperably coupled to the first turbine assembly; and a turbine framepositioned in serial flow arrangement downstream of the first turbineassembly, and wherein the first location at the turbine section is atthe turbine frame.
 9. The gas turbine engine of claim 1, the gas turbineengine comprising: a fluid system configured to provide a flow of heattransfer fluid in thermal communication with the flow of air via thefirst heat exchanger.
 10. The gas turbine engine of claim 1, wherein thefirst conduit comprises: a plurality of inlet manifolds configured toreceive the flow of air from a plurality of circumferential compressorlocations at the compressor section; and a collector configured toreceive the flow of air from the plurality of inlet manifolds, andwherein the collector is fluidly coupled to the first heat exchanger toprovide the flow of air in thermal communication to the first heatexchanger.
 11. The gas turbine engine of claim 1, wherein the engine isconfigured to operate between approximately 75% and approximately 90% ofan overall power output of the engine.
 12. The gas turbine engine ofclaim 11, wherein the first conduit is configured to extract the flow ofair from the compressor section at a compressor location having anairflow pressure between approximately 20 pounds per square inch andapproximately 60 pounds per square inch while operating the enginebetween approximately 75% and approximately 90% of the overall poweroutput.
 13. The gas turbine engine of claim 1, wherein the engine isconfigured to operate at an operating condition corresponding to betweenapproximately 55% and approximately 75% of an operating envelope;wherein the first conduit is configured to extract the flow of air fromthe compressor section at a compressor location having an airflowpressure between approximately 20 pounds per square inch andapproximately 60 pounds per square inch during the operating conditioncorresponding to between approximately 55% and approximately 75% of theoperating envelope.
 14. The gas turbine engine of claim 1, wherein thecompressor section comprises a low speed compressor and a high speedcompressor, and wherein the turbine section comprises a low speedturbine, a high speed turbine, and a turbine frame positioned betweenthe low speed turbine and the high speed turbine, and wherein the firstlocation at the turbine section is at the turbine frame, and wherein thefirst conduit extends in fluid communication from the high speedcompressor to the turbine frame.
 15. The gas turbine engine of claim 1,wherein the first conduit is configured as a fixed area flowpath fromthe compressor section to the turbine section, and wherein the flowcontrol device defines a variable area flowpath at the second conduit.16. The gas turbine engine of claim 15, wherein the first conduit isconfigured continuously flow the extracted flow of air to the firstlocation at the turbine section through an operating condition of theengine, and wherein the flow control is configured to provide a variableflow to the second location at the turbine section.
 17. The gas turbineengine of claim 1, wherein the gas turbine engine further comprises aclearance control system positioned at the second location at theturbine section, wherein the clearance control system is configured toselectively altering a tip clearance based on the portion of the flow ofair extracted from the first conduit to the second conduit.
 18. Anairflow delivery system for a gas turbine engine, the gas turbine enginecomprising a compressor section and a turbine section in serial flowarrangement, the airflow delivery system comprising: a first conduitextended from the compressor section to the turbine section, the firstconduit in fluid communication with the compressor section and theturbine section to communicate a flow of air from the compressor sectionto a first location at the turbine section; a first heat exchangerpositioned in thermal communication with the flow of air through thefirst conduit; a second conduit in fluid communication with the firstconduit at a location downstream of the first heat exchanger and influid communication with a second location at the turbine section,wherein the second conduit is configured to receive a portion of theflow of air from the first conduit, and wherein the second conduit isconfigured to deliver only the portion of the flow of air to the secondlocation at the turbine section; and a flow control device positioned inflow communication with the second conduit configured to modulate anamount of the flow of air from the first conduit through the secondconduit.